U.S. patent application number 14/819792 was filed with the patent office on 2017-02-09 for method for the joining of wings or control surfaces to an airplane fuselage.
The applicant listed for this patent is The Boeing Company. Invention is credited to Jonathan D. Embler, Thomas R. Pinney, Richard P. Topf.
Application Number | 20170036751 14/819792 |
Document ID | / |
Family ID | 58052354 |
Filed Date | 2017-02-09 |
United States Patent
Application |
20170036751 |
Kind Code |
A1 |
Topf; Richard P. ; et
al. |
February 9, 2017 |
METHOD FOR THE JOINING OF WINGS OR CONTROL SURFACES TO AN AIRPLANE
FUSELAGE
Abstract
Methods of joining a first aero structure part with a second
aero structure part are disclosed, for example, methods of joining
an airplane wing or control surface to an airplane fuselage or
joining fins to a rocket body. The method comprises aligning a
plurality of connection elements in a linear array within the
second aero structure part, wherein the connection elements
comprise a plurality of flexible connection elements and at least
one rigid connection element, and attaching the first aero
structure part to the second aero structure part at a plurality of
connection points with the plurality of connection elements. The
second aero structure part expands linearly at a greater rate when
exposed to heat than the first aero structure part resulting in a
difference in linear distance between the first aero structure part
and the second aero structure part and the flexible connection
elements are configured to flex to accommodate for this linear
distance difference.
Inventors: |
Topf; Richard P.; (Chicago,
IL) ; Embler; Jonathan D.; (Chicago, IL) ;
Pinney; Thomas R.; (Chicago, IL) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
The Boeing Company |
Chicago |
IL |
US |
|
|
Family ID: |
58052354 |
Appl. No.: |
14/819792 |
Filed: |
August 6, 2015 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B64C 1/26 20130101; Y02T
50/44 20130101; B64C 1/38 20130101; B64C 9/02 20130101; Y02T 50/40
20130101; B64C 30/00 20130101 |
International
Class: |
B64C 1/38 20060101
B64C001/38; B64C 9/02 20060101 B64C009/02; B64C 30/00 20060101
B64C030/00; B64C 1/26 20060101 B64C001/26 |
Claims
1. A method of joining a first aero structure part with a second
aero structure part comprising aligning a plurality of connection
elements in a linear array within the second aero structure part,
wherein the connection elements comprise a plurality of flexible
connection elements and at least one rigid connection element, and
attaching the first aero structure part to the second aero
structure part at a plurality of connection points with the
plurality of connection elements, wherein the second aero structure
part expands linearly at a greater rate when exposed to heat than
the first aero structure part resulting in a difference in linear
distance between the first aero structure part and the second aero
structure part, and wherein the flexible connection elements are
configured to flex to accommodate for this linear distance
difference.
2. The method of claim 1, wherein the flexible connection elements
are flexible plates.
3. The method of claim 2, wherein the flexible plates each comprise
at least one stack of plates.
4. The method of claim 2, wherein the flexible plates have an "A"
shape profile.
5. The method of claim 1, wherein the first aero structure part is
an airplane fuselage.
6. The method of claim 1, wherein the second aero structure part is
an airplane wing.
7. The method of claim 1, wherein the second aero structure part is
a control surface.
8. The method of claim 1, further comprising positioning the rigid
connection element such that the amount of flex in the plurality of
flexible connection elements is equalized on both sides of the
rigid connection element.
9. The method of claim 6, further comprising aligning the plurality
of flexible connection elements such that the thinnest dimension of
the flexible connection element is in a chordwise direction of the
airplane wing.
10. The method of claim 1, wherein the flexible connection elements
are flexible in a chordwise direction and rigid in other load
directions.
11. The method of claim 3, wherein the flexible plates comprise at
least two stacks of flexible plates and wherein the flexible plates
comprise a flexible spacer between the stacks of plates configured
to prevent buckling of the flexible plates.
12. The method of claim 1, wherein the flexible connection elements
comprise metallic material.
13. The method of claim 1, wherein the flexible connection elements
comprise ceramic material.
14. A flexible connection element for use in the method of claim 1,
wherein the flexible connection element comprises a plurality of
stacked plates, and wherein the flexible connection element is
A-shaped.
15. A method of joining an airplane wing with an airplane fuselage
comprising aligning a plurality of flexible plates and one rigid
connection element in the airplane wing, such that the thinnest
part of each flexible plate is aligned with a chordwise direction
of the airplane wing and the length of each flexible plate is
aligned with a spanwise direction of the airplane wing, and
attaching the wing to the fuselage with the flexible plates and the
rigid connection element at a plurality of connection points,
wherein the rigid connection element connects the wing to fuselage
at a fixed location, wherein the flexible plates flex to
accommodate chordwise expansion of the airplane wing, and wherein
the flexible plates resist vertical shear and bending moment.
16. The method of claim 15, further comprising positioning the
rigid connection element such that the amount of flex in the
plurality of flexible connection element is equalized on both sides
of the rigid connection element.
17. The method of claim 15, wherein the flexible plates each
comprise at least one stack of plates.
18. The method of claim 17, wherein the flexible plates comprise at
least two stacks of plates and wherein the flexible plates comprise
a flexible spacer between the stacks of plates configured to
prevent buckling of the flexible plates.
19. The method of claim 15, wherein the airplane wing comprises
ceramic matrix composite material and wherein the airplane fuselage
comprises metallic materials.
20. A flexible connection element for use in the method of claim
15, wherein the flexible connection element comprises a plurality
of stacked plates, and wherein the flexible connection element is
A-shaped.
Description
TECHNICAL FIELD
[0001] The present disclosure relates generally to methods of
joining two aero structures, for example, joining wings or control
surface structures to an airplane fuselage or joining fins to a
rocket body. Specifically, this disclosure relates to using
flexible connection elements to join aero structures with different
thermal linear expansion rates, different thermal environments, or
different thermal management approaches.
BACKGROUND
[0002] Vehicles that travel at hypersonic speeds need an airframe
designed to withstand thermal loads as well as structural loads.
This is especially true for airframes designed with hot structure,
whereby a portion of the vehicle structure is allowed to get hot,
as opposed to a more traditional approach of using thermal
protection materials on the vehicle surface. Structural
configurations for wing-to-body joints that are typically used for
aircraft are not appropriate for a vehicle with a hot structure
wing. This is because the rigid connection of a traditional joint
cannot accommodate the strain induced by thermal expansion,
especially along the chord of the wing.
[0003] In most aero structures, one or more primary load bearing
members called spars create just a few paths to react to primary
bending and shear loads. The spars may pass through the fuselage,
under or over the fuselage, or connect directly to the fuselage.
Regardless of position, a rigid connection is used to transfer wing
loads into the fuselage. Typically, hypersonic vehicles are
designed with parasitic thermal protection systems (TPS) on the
skins of the vehicle that minimize internal structure temperature,
thus enabling the traditional structural approach. TPS adds
significant weight to the vehicle. It also adds cross sectional
area to the vehicle and thickness to the wings in particular, which
add significant drag forces at hypersonic speeds. A hot structure
wing, whereby the wing is allowed to get hot from aerodynamic
heating, would thus be much more efficient due to lower weight and
reduced thickness. However, the thermal growth of the wing relative
to the fuselage at extreme temperatures would overstress a
traditional rigid connection, making it infeasible.
[0004] Thus, there is a need for a method of joining hot structure
wings or control surfaces to the fuselage without overstressing the
connection when relative thermal growth exists between structural
members.
SUMMARY
[0005] According to an example embodiment, a method of joining a
first aero structure part with a second aero structure part is
provided. The method comprises aligning a plurality of connection
elements in a linear array within the second aero structure part,
wherein the connection elements comprise a plurality of flexible
connection elements and at least one rigid connection element. The
method further comprises attaching the first aero structure part to
the second aero structure part at a plurality of connection points
with the plurality of connection elements. The second aero
structure part expands linearly at a greater rate when exposed to
heat than the first aero structure part resulting in a difference
in linear distance between the first aero structure part and the
second aero structure part, and wherein the flexible connection
elements are configured to flex to accommodate for this linear
distance difference.
[0006] According to another example embodiment, a method of joining
an airplane wing with an airplane fuselage is provided. The method
comprises aligning a plurality of flexible plates and one rigid
connection element in the airplane wing, such that the thinnest
part of each flexible plate is aligned with a chordwise direction
of the airplane wing and the length of each flexible plate is
aligned with a spanwise direction of the airplane wing. The method
further comprises attaching the wing to the fuselage with the
flexible plates and the rigid connection element at a plurality of
connection points, wherein the rigid connection element connects
the wing to fuselage at a fixed location. The flexible plates flex
to accommodate chordwise expansion of the airplane wing, and
wherein the flexible plates resist vertical shear and bending
moment.
[0007] The features, functions, and advantages can be achieved
independently in various embodiments of the present disclosure or
may be combined in yet other embodiments in which further details
can be seen with reference to the following description and
drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] The novel features believed characteristic of the example
embodiments are set forth in the appended claims. The example
embodiments, however, as well as a preferred mode of use, further
objectives and descriptions thereof, will best be understood by
reference to the following detailed description of an example
embodiment of the present disclosure when read in conjunction with
the accompanying drawings, wherein:
[0009] FIG. 1 is a diagrammic representation of a portion of an
airplane showing a fuselage and a wing joined together using a
method of the present disclosure;
[0010] FIG. 2 is a diagrammic representation of a portion of an
airplane showing a fuselage and a wing joined together with
flexible connection elements using a method of the present
disclosure;
[0011] FIGS. 3a and 3b are diagrammic representations of a flexible
connection element of the present disclosure;
[0012] FIG. 4 is a diagrammic representation of a block diagram of
an aircraft manufacturing and service method in accordance with an
example embodiment; and
[0013] FIG. 5 is a diagrammic representation of a block diagram of
an aircraft in which an example embodiment may be implemented.
DETAILED DESCRIPTION
[0014] Disclosed embodiments will now be described more fully
hereinafter with reference to the accompanying drawings, in which
some, but not all of the disclosed embodiments are shown. Indeed,
several different embodiments may be provided and should not be
construed as limited to the embodiments set forth herein. Rather,
these embodiments are provided so that this disclosure will be
thorough and complete and will fully convey the scope of the
disclosure to those skilled in the art.
[0015] This disclosure seeks to provide a solution to problems that
can arise when designing an attachment method for wings and/or
control surfaces to aero vehicle structures when there is a large
difference in thermal growth due to aerodynamic heating at high
speeds. While material systems have been developed for dealing with
high and low temperatures within major structural parts, joining
these can be problematic. This disclosure describes a method using
multiple flexible connection elements, such as flexible plates,
along a line to form the connection between major structures, such
as wings and fuselage, or fins and rocket bodies. The disclosed
method could also be used to join parts in other fields, wherein
the two parts have different thermal growth rates.
[0016] This disclosure provides a method for joining wings and or
control surfaces to a payload cabin (fuselage) that allows for
differential thermal growth between joint interfaces. The method
uses a plurality of connection elements with flexibility in the
chordwise direction, allowing for thermal growth of the wing. The
geometry of the flexible connection elements is such that there is
significantly more stiffness in the vertical and spanwise
directions. A plurality of such flexible connection elements are
used, each with a reduced overall load that can be reacted without
the need of significant beam type spars. The loads are distributed
among each location with an appropriately size connection element
to the fuselage shell as its load path. By making these connection
elements from relatively thin, flexible plates, changes in the
length of the fuselage relative to the wing on vehicles with high
aerodynamic heating can be accommodated through flexure of the thin
plates. Metal alloys, typically with high nickel content could be
used for the plate elements. These metal alloys have high
temperature resistance whilst also having a much higher elastic and
tensile modulus than other high temperature materials such as
ceramics. The coefficient of thermal expansion of the aero parts
will depend on the material. For example, a carbon composite will
have a low coefficient of thermal expansion, while a metal will
have a higher coefficient of thermal expansion. In some cases it
may be beneficial to have many thin plate elements stacked together
to improve buckling performance while maintaining flexibility. This
could be accomplished by layering joint elements with ceramic
fabric, felt or another material between them.
[0017] FIGS. 1 and 2 show a portion of an airplane 100 with a first
aero structure part 102 and a second aero structure part 104. The
first aero structure part 102 and the second aero structure part
104 are joined at connection points 106. A method of joining the
first aero structure part 102 with the second aero structure part
104 is provided. The second aero structure part 104 may expand
linearly at a greater rate when exposed to heat than the first aero
structure part 102, resulting in a difference in linear distance
between the first aero structure part 102 and the second aero
structure part 104. In an example embodiment, the first aero
structure part 102 may be an airplane fuselage and the second aero
structure part 104 may be an airplane wing, as shown in FIGS. 1 and
2. The first aero structure part/airplane fuselage 102 may comprise
metallic materials and the second aero structure part/airplane wing
104 may comprise ceramic matrix composite material. In other
embodiments, the second aero structure part may be a control
surface. In yet other embodiments, the first aero structure part
may be a rocket body and the second aero structure part may be a
fin.
[0018] The method may comprise aligning a plurality of connection
elements in a linear array within the second aero structure part
104. The connection elements may comprise a plurality of flexible
connection elements 108 and at least one rigid connection element
110. The rigid connection element 110 may be located in between the
flexible connection elements 108. For example the rigid connection
element 110 may be located near the middle of the flexible
connection elements 108. In other embodiments, the rigid connection
element 110 may be located at an end of the linear array. In yet
other embodiments, the rigid connection element 110 may be offset
from the flexible connection elements 108. The rigid connection
element 110 may include wires connecting the first aero structure
part 102 with the second aero structure part 104. The method may
further comprise attaching the first aero structure part 102 to the
second aero structure part 104 at a plurality of connection points
106 with the plurality of connection elements. The flexible
connection elements 108 may be configured to flex to accommodate
for the difference in linear thermal growth between the first aero
structure part 102 and the second aero structure part 104.
[0019] An example of a flexible connection element 108 of the
present disclosure is shown in FIGS. 3a and 3b. The flexible
connection elements 108 may be flexible plates. The flexible plates
may each comprise at least one stack of plates 116. In an example
embodiment, the flexible plates may comprise at least two stacks of
plates 116 and a flexible spacer 118 between the stacks of plates
116 configured to prevent buckling of the flexible plates and to
provide out-of-plane support (i.e. stabilizing the stacks such that
more bending moment is required to buckle the stacks). The flexible
plates may have an "A" shape profile, wherein the wider end 120
(the bottom of the "A") is attached to the first aero structure
part 102 and the narrower end 122 (the top of the "A") is attached
to the second aero structure part 104. The flexible connection
elements 108 may connect the first aero structure part 102 to the
second aero structure part 104 with fastener, for example, a clevis
and pin joint. In other embodiments, the flexible plates may have
an "X" shaped profile or be a truss-like member. Other shapes could
also be used depending on the load requirements, wing thickness,
and other factors as long as such shapes are flexible in the
chordwise direction. The wider end 120 and the narrower end 122 may
include a plurality of holes 124, which assist in attaching the
first aero structure part 102 to the second aero structure part
104. The flexible connection elements 108 may comprise a metallic
material, for example, nickel or titanium alloys. In considering
which metal alloys to use, elastic modulus (stiffness), fatigue
resistance, strength and weight at high temperatures, and other
properties are considered. In other embodiments, the flexible
connection elements 108 may comprise a ceramic material. The
flexible spacer(s) 118 may comprise ceramic felt, ceramic fabric,
metallic mesh, metallic sheet, polymer material (e.g., silicone),
or any other material known in the art. The flexible connection
elements 108 may be spaced at equal distances from each other. In
other embodiments, the flexible connection elements 108 may be
spaced at unequal distances from each other. The distance between
the flexible connection elements 108 and the number of flexible
connection elements 108 is determined by aircraft specific
loads.
[0020] The rigid connection element 110 may be positioned such that
the amount of flex in the plurality of flexible connection elements
108 is equalized on both sides of the rigid connection element 110.
The rigid connection element 110 may be made of any material that
can withstand high temperatures. The rigid connection element 110
may be arranged in a manner similar to a traditional wing-to-body
joint arrangement, for example, as a bolted joint connection.
[0021] The method may further comprise aligning the plurality of
flexible connection elements 108 such that the thinnest dimension
of each flexible connection element 108 is in a chordwise direction
112 and the length of each flexible connection element 108 is
aligned with a spanwise direction 114 of the airplane wing 104. The
thinnest dimension of each flexible connection element 108 is
defined by the layers of the flexible plates 116 and flexible
spacers 118. The flexible connection elements 108 may be flexible
in the chordwise direction 112 and rigid in other load
directions.
[0022] Examples of the disclosure may be described in the context
of an aircraft manufacturing and service method 200 as shown in
FIG. 4 and an aircraft 250 as shown in FIG. 5. Turning first to
FIG. 4, an illustration of an aircraft manufacturing and service
method is depicted in accordance with an example. During
pre-production, exemplary method 200 may include specification and
design step 202 of the aircraft 250 in FIG. 5 and material
procurement step 204.
[0023] During production, component and subassembly manufacturing
step 206 and system integration step 208 of the aircraft 250 in
FIG. 5 takes place. For example, the first aero structure part 102
may be attached to the second aero structure part 104 during step
206. Thereafter, the aircraft 250 in FIG. 5 may go through
certification and delivery step 210 in order to be placed in
service step 212. While in service by a customer, the aircraft 250
in FIG. 5 is scheduled for routine maintenance and service step 214
(which may also include modification, reconfiguration,
refurbishment, and so on).
[0024] Each of the processes of method 200 may be performed or
carried out by a system integrator, a third party, and/or an
operator (e.g., a customer). For the purposes of this description,
a system integrator may include without limitation any number of
aircraft manufacturers and major-system subcontractors; a third
party may include without limitation any number of venders,
subcontractors, and suppliers; and an operator may be an airline,
leasing company, military entity, service organization, and so
on.
[0025] With reference now to FIG. 5, an illustration of an aircraft
250 is depicted in which an example may be implemented. In this
example the aircraft 250 produced by exemplary method 200 in FIG. 4
and may include an airframe 252 with a plurality of high-level
systems 254 and an interior 256. Examples of high-level systems 254
include one or more of a propulsion system 218, an electrical
system 220, a hydraulic system 222, and an environmental system
224. In an example, the method of the present disclosure may be
used to manufacture the air. Any number of other systems may be
included. Although an aerospace example is shown, the principles of
the disclosure may be applied to other industries, such as the
automotive industry.
[0026] Systems and methods embodied herein may be employed during
any one or more of the stages of the production and service method
200 in FIG. 4. For example, components or subassemblies
corresponding to production step 206 may be fabricated or
manufactured in a manner similar to components or subassemblies
produced while the aircraft 250 is in service. Also, one or more
apparatus examples, method examples, or a combination thereof may
be utilized during the production step 206 and certification and
delivery step 210, for example, by expediting assembly of or
reducing the cost of an aircraft 250. Similarly, one or more of
apparatus examples, method examples, or a combination thereof may
be utilized while the aircraft 250 is in service, for example and
without limitation, to maintenance and service 214.
EXAMPLES
[0027] In an illustrative example, the hypersonic vehicle in which
the wing is to be attached to the fuselage is a reusable system
with both take-off and landing capacity. The wing is approximately
11-17 inches thick with a thin and sharp leading edge (about 7.5
inches thick) and fixed trailing edges. The wing is made out of
ceramic matrix composite type materials and the wing does not
include TPS. The temperature of the wing during flight is about
1400 degrees Fahrenheit. The fuselage is made out of a metallic and
organic composite material (e.g., steel, carbon/bismaleimide,
aluminum) and is protected by TPS. The temperature of the fuselage
during flight is about 400 degrees Fahrenheit. The coefficient of
thermal expansion for the wing is between 1.5 and 4.5 ppm/degree
Fahrenheit. The coefficient of thermal expansion for the fuselage
is 1.5 ppm/degree Fahrenheit. Total shear is 240 k and moment is
25.times.10.sup.6 in-lb. The length of the wing at the
fuselage/wing interface is about 428 inches and the distance
between connection points is about 13.9 inches. In this example,
the length of the wing at the fuselage/wing interface expands
between 0.90 inches and 2.70 inches and the distance between
connection points expands between 0.03 inches and 0.09 inches. The
flexible connection elements accommodate for this expansion by
flexing in the chordwise direction.
[0028] The foregoing description of the specific embodiments will
reveal the general nature of the disclosure so others can, by
applying current knowledge, readily modify and/or adapt for various
applications such specific embodiments without departing from the
generic concept, and therefore such adaptations and modifications
are intended to be comprehended within the meaning and range of
equivalents of the disclosed embodiments. It is to be understood
that the phraseology or terminology herein is for the purpose of
description and not of limitation.
* * * * *